1. Field of the Invention
The invention relates to a gas turbine engine rotor blading sealing device permitting a real time adjustment to be made of clearances between the rotor blade tips and the stationary surrounding structure. The invention also defines a method for determining if the sealing device conforms to its intended function.
2. Brief Description of the Prior Art
It is important to reduce fluid leaks between the rotating and fixed parts of a gas turbine engine, particularly in relation to the turbine, since they affect the efficiency, the maximum power and the resistance to hunting of the engine.
In order to reduce these clearances and, correspondingly, the leaks, whether the engine is operating in a stabilized or transient mode, it is necessary to meet certain conditions, some of which are incompatible with the others. These conditions involve the concentricity of the tips of the blades, or of their peripheral roots, and the concentricity of the sealing device with the rotational axis of the gas turbine engine; the undeformability of the device (i.e. its ability to resist deformation from a circular cross section); the increase or decrease in the radius of the sealing device with the increase or decrease in the radius of the blade tips or their peripheral roots, under the effect of centrifugal and thermal expansions, whether stabilized or transient.
It is a relatively simple matter to insure that the tip of the rotating parts (blade tips or peripheral roots) describes a surface of revolution. This can be accomplished by machining the blades or the roots to a predetermined length by grinding on a rotating blades wheel. However, it is much more difficult to provide a shroud surrounding the rotating parts with a form of revolution capable of withstanding the different operating conditions. Particular factors which render this difficult are the thermal deformations caused by changes in the operating temperature of the engine and inertial forces (load factors in the Z or Y direction in particular) caused by the variations in operating rpms. It is also necessary that the shroud be impervious, that it remains perfectly centered on the axis of rotation of the gas turbine engine, and that it resists deformation and maintain its circular cross section. These factors make it obligatory that the shroud be designed as a monolithic structure having adequate inertia, or as a more complicated system having means to ensure the concentricity of the supports of the shroud sectors in relation to the engine axis and the complete absence of any ovalness.
It these conditions are not met and the shroud assumes a degree of ovalness or eccentricity such that the maximum distance inward between a circle of the same developed length and the interior part of the oval shaped (or eccentric) shroud closest to the blade tips (or to the segment supports) with maximum ovalness (or eccentricity) is designated by "a", is necessary to incorporate into the design:
a clearance greater than or equal to "a" under design operating conditions; or, PA1 a packing of minimal thickness, equal to "a" that will be locally removed by abrasion upon the appearance of the maximum ovalness (or eccentricity) "a", thus forming a leak area through a clearance that is locally equal to "a". In all likelihood, this leak area will extend around the entire periphery of the shroud since the ovalness (or eccentricity) can occur at different times along different axes. PA1 the radial dimensions of the shroud or the shroud sectors maintain a small, positive clearance between the blade tips (or their peripheral roots) and the shroud or shroud sectors.
Means are known to have the casing supporting the shroud well centered with respect to the axis of the gas turbine engine, as described by the applicants in their patent application Ser. No. 81.20719 filed on Nov. 5, 1981. Means are also known to center the shroud in the casing and to give it sufficient inertia such that its deformation into an oval shape is practically negligible. However, ideal sealing shrouds should ensure that, under all operating conditions, particularly the transient conditions from idle speed to full throttle and vice versa,
In the following explanation, for reasons of simplification, reference shall be made to a structure in which the rotor blades do not have peripheral roots, but it should be understood that the term "blade tip" includes the radial tip of the peripheral roots of the blades when the blades have such peripheral roots.
The first difficulty that is encountered in designing an effective sealing shroud relates to the fact that the dimension of the shroud must be adjusted in order to prevent or minimize leaks through all modes of the engine operation from rest, through idling speed to maximum speed. If r.sub.o is the sum of the radius of the turbine wheel plus the adjusted length of the blades at rest, the radial position of the blade tip for a stabilized idling speed (assuming for the sake of simplification that the blade tips sweep a cylinder and not some other form of revolution): ##EQU1## and, for the maximum stabilized mode, replacing subscript r for idling speed by subscript m for maximum EQU r.sub.m =r.sub.o +dcd.sub.m +dca.sub.m +dtd.sub.m +dta.sub.m
and doing the same for any intermediate mode (indicated by subscript i) EQU r.sub.i =r.sub.o +dcd.sub.i +dca.sub.i +dtd.sub.i +dta.sub.i
It is for these respective radii, r.sub.r, r.sub.m and r.sub.i that it is necessary to adjust the inside dimension of the shroud to maintain the clearance as small as possible.
If the shroud is a monolithic structure having sufficient inertia to maintain its concentricity about the engine axis and to maintain its circular configuration, the simplest way of causing the inside radius of the shroud or shroud sectors to vary, is to vary its temperature. This may be accomplished by selecting the shroud material to have a coefficient of thermal expansion .alpha. such that a very small positive clearance may be maintained under different stabilized conditions by directing heated air taken from an appropriate stage of the gas turbine engine compressor onto the shroud structure.
Many ways of accomplishing this are known, a typical examples of which is shown in French Patent 2,467,292. None of the prior art, however, has completely resolved the problem even when the operating mode of the engine is stabilized. These known means simple encompass an air distribution mechanism which may be adjusted for flow and/or temperature variations to direct heated air onto the shroud structure in order to vary its temperature thereby maintaining a small clearance.
While these systems can theoretically be made, they are extremely complicated and their reliability in everyday usage is quite problematical. In some cases, breakdown of the air distribution mechanism is capable of causing significant damage to the turbine assembly and/or the shroud. A more serious defect of these known systems is that they do not account for factors which will maintain the appropriate blade tip clearances during transition from one operational mode to another. To be completely effective, a shroud sealing system must have a response which adjusts to the response of the radial displacement of the blade tip due to the expansion or contraction of the rotor during the transition of the engine from one operation mode to another.
The use of these prior art devices during rapid deceleration of the gas turbine engine, such as that often required of the engine in aviation usage, would result in an increase in the clearance until the shroud structure has had a chance to stabilize at a new, lower temperature. The periods of rapid deceleration and acceleration are on the order of six seconds in typical aviation usage. During this transient period, there is a radial displacement of the blade tips inwardly due to the decrease in centrifugal forces acting on the blades and the wheel. This radial displacement is equal to (dcd.sub.m -dcd.sub.r)+(dca.sub.m -dca.sub.r). Obviously, if the radius of the shroud carrying the sealing sectors has not varied, the clearance between it and the blade tips would increase until the temperature of the shroud could be stabilized at the new operating mode.
The time period of six seconds will be used in describing the transient modes for both acceleration and deceleration of the rotating parts of the engine, but obviously the invention encompasses longer or shorter acceleration or deceleration times.
In approximately the same period of time that it takes for the engine to decelerate as noted above, the gases driving the turbine wheel, the fluid cooling the wheel, and the fluid cooling the interior of the turbine blades enable the wheel and blades to reach their corresponding idling speed temperatures. The following analysis will disregard the heating of the ventilating air as a result of the thermal inertia of those parts in contact with these fluids, such as piping, enclosures, etc. This results in a thermal contraction of the blade (dta.sub.m -dta.sub.r) that is added to the contraction (dcd.sub.m -dcd.sub.r +dca.sub.m -dca.sub.r) due to the reduction in the centrifugal forces. It must be pointed out, however, that responses, as a function of time, of the respective thermal expansions and contractions of the blade and wheel, indeed even the different parts of the wheel, are very different. To simplify the analysis, the following will only be concerned with the time that it takes one of these members to acquire a thermal expansion (or contraction) equal to 98% of its final expansion (or contraction) to a completely stabilized mode. This will be referred to as the "98% response time" or simply the "response time".
A value that is characteristic of this response time for a modern turbine blade design is on the order of a few seconds. Contrasted to this, the turbine wheel, because of its considerably larger thermal inertia takes on the order of fifty times longer (and in some cases even longer) to reach its quasi stabilized temperature. The relatively thin rim part of the turbine wheel (which is still thicker than the turbine blades) will reheat relatively quickly during acceleration since it is the direct recipient of the flow of heat from the turbine blades. The central portion of the turbine wheel which is generally much thicker, takes longer to reheat and, thus its response time is greater than the rim or the blades. The response time for the turbine wheel will be defined as the time it takes the wheel to reach its 98% thermal expansion (or contraction) value in a completely stabilized mode.
The known prior art devices do not enable the shroud structure to make a "real time" adjustment to centrifugal and thermal expansion of the turbine blade wheel. If the shroud structure carrying the seal sectors is designed to have the same response time as the wheel (by properly choosing its dimensions and heat insulation, for example) during the sudden deceleration noted above, the clearance between the blade tips and the shroud seal will be increased by a dimension equal to: EQU dcd.sub.m +dca.sub.m +dta.sub.m -dcd.sub.r -dca.sub.r -dta.sub.r
This increase in clearance is not critical since it applies only to a transient operation during which time there is no wear of the sealing material. After thermal contraction of the turbine wheel occurs, the stabilized idling speed radius is equal to: EQU r.sub.o +dcd.sub.m -dcd.sub.r +dca.sub.m -dca.sub.r +dta.sub.m -dta.sub.r +dtd.sub.m -dtd.sub.r
By using the prior art devices, it is possible to provide a very small positive clearance between the blade tips and the shroud during the stabilized idling speed and the stabilized maximum speed mode by a judicious choice of the shroud material, particularly with respect to its coefficient of thermal expansion .alpha.. However, the clearance would become a negative value during sudden acceleration of the turbine blade wheel thereby causing breakage of the blades or significant wear of the abradable seal material. During the rapid acceleration (within a period of approximately seven seconds) the radius of the blades will be equal: EQU dcd.sub.m -dcd.sub.r +dca.sub.m -dca.sub.r +dta.sub.m -dta.sub.r
This increase can be on the order of 1.5 millemeters or greater. An increase of this magnitude at the same time that the radius of the shroud has not begun to expand will result in severe damage to the engine.
Even if the turbine blade tips are not damaged, they will abrade the shroud sealing material, resulting in an enlarged clearance after the shroud has expanded and stabilized. Once the seal material has been abraded away, the enlarged clearances will remain under other operating conditions.
If the shroud structure has been designed in conjunction with the characteristics of its ventilation air to have the same response time as the engine itself during acceleration and/or deceleration approximately equal to the centrifugal expansion/contraction response time of the wheel and blade plus the thermal expansion/contraction of the blade, it is by going from full stabilized throttle that a sudden reduction would result in negative clearance, to the reduced value dtd.sub.m -dtd.sub.r for the clearance at the outset for the stabilized maximum. This negative clearance would abrade a layer of the seal element (on the order of 2.5 millemeters) resulting in a clearance that would remain in subsequent operation modes.
It is possible to refine the air distribution system further such that is responds approximately in real time to the double response time curve for the radius of the blade tips as shown in FIG. 1 (acceleration being above the time axis and deceleration being below the time axis). However, this results in an increase in complexity, mass, and cost, and a decrease in the reliability of the system. Furthermore, the piping must be designed for the maximum ventilation flow which would influence the temperature of the shroud structure during the first phase of starting and the temperature of the wheel and turbine blades.
Attempts to solve the problems posed by the necessity of the shroud structure and sealing sectors having to respond to the double response time curve, shown in FIG. 1, are known and typical examples are illustrated in French Patents 2,450,344 and 2,450,345. The arrangement provided for in these patents are applicable only to low-powered gas turbine engines with reverse flow combustion chambers. The principles set forth in these examples of the prior art could possibly be adapted to direct flow combustion chambers for high-powered gas turbine engines, but they would be extremely expensive and unrealistically complicated.
The solution to the problems set forth in the French patents noted above involves the use of a flexible sleeve, one that can deform when subjected to stresses. This solution presents the risk and disadvantages of introducing defects in the concentricity and ovalness of the shroud structure, particularly under the effect of load factors encountered in flight. As previously discussed, the exact concentricity of the shroud structure with respect to the axis of rotation of the engine and the absence of any ovalness of the shroud structure and sealing sectors are primary considerations for any successful sealing system. The systems set forth in the French patents noted above do not satisfactorily meet these criteria. It is also noted that due to the considerable hyperstatic stresses brought into play by the seizures in a segmented ring as set forth in French Patent 2,450,345, the least heterogeneity in temperature or inertia in the peripheral direction will cause significant deformations of the segmented ring.